This invention generally relates to energy-exchange systems, and more particularly to energy-exchange systems utilizing relatively small-diameter, or "capillary", tubes incorporated beneath the skin of an aircraft.
Aircraft desirably operate in all weather conditions, including very low temperature conditions in which ice may tend to accrete upon an aircraft surface. This ice accretion occurs particularly along the leading edges of the aircraft, such as along the forward portion of the aircraft wing. It is desirable to either remove the ice following its accretion (de-icing) or, more preferably, to prevent the ice from accreting (anti-icing), because it may interfere with the proper aerodynamic function of the aircraft or increase the aircraft's weight. It is also preferable that ice be prevented from accreting because the ice, during its removal, may damage aircraft components with which it would collide downstream.
Several alternative approaches have been proposed to provide thermal anti-icing systems for aircraft, including that disclosed in U.S. Pat. No. 5,011,098 issued to McClarin, et al, in which portions of the aircraft surface for which anti-icing is desired are comprised of an integrated skin having passageways formed therein in which heated fluid may be passed. The anti-icing system of the McClarin patent requires integrated inner and outer skins, however, so as to limit the flexibility of the designer in determining the aircraft's structure. Additionally, the large cavity between the inner and outer skins would make effective use of a liquid working fluid difficult due to the additional weight of the liquid.
A structural cooling unit has been disclosed by Niggemann in U.S. Pat. No. 4,786,015 in which an aircraft's leading edges and nose cones have a helically wound metallic tube formed within their exterior surfaces to form a single flow path incorporated therein. Such a cooling unit as that described by Niggemann would both be difficult to fabricate and result in variable cooling. Because only a single coolant path is provided, the coolant would gradually absorb heat from the aircraft's outer surface, and thus have its ability to cool the portion of aircraft's surface located downstream impaired.
U.S. Pat. No. 2,645,435 issued to Pouit discloses another energy-exchange system for an aircraft's leading edge in which passageways are formed by internal reinforcements within a double-wall structure in order to transport fluids therein. Such a structure requires double walls for the aircraft's leading edge, so as to restrict the designer's flexibility in choosing alternative structural designs.
An aircraft anti-icing plenum is disclosed by Cook, et al in U.S. Pat. No. 3,933,327 in which the leading edge of a jet engine nacelle is fabricated to prevent formation of ice thereon. Such a design as that illustrated in U.S. Pat. No. 3,933,327 requires a double-wall construction for the leading edge, however, and would be unsuitable for use with a liquid working fluid, as the fluid is exhausted to atmosphere following its traversal through the energy-exchange system. The exhaust of a liquid working fluid at the leading edge of an aircraft engine may cause difficulties with aircraft components which the exhausted liquid would contact downstream. Additionally, the large plenum filled with a liquid would weigh a substantial amount and cause increased weight for the aircraft to carry.
General Electric Company, in a report entitled Evaluation of Capillary Reinforced Composites submitted to the National Astronautics and Space Administration (NASA) under contract number NASA-CR175061 for the period from September 1984 through September 1985, as well as an article by Ciardullo, et al of General Electric Company entitled Evaluation of Capillary Reinforced Composites for Anti-Icing submitted at the AIAA 25th Aerospace Sciences meeting held Jan. 12-15, 1987, disclosed the use of glass capillary tubes for exchanging energy with composite skins. The use of glass tubes disclosed by the General Electric reports suffers from several deficiencies, including the susceptibility of glass to long-term stress corrosion and contamination corrosion and the difficulty in bending glass tubes to hold a relatively small radius of curvature. An additional difficulty with the General Electric design occurs in the connection of the glass capillary tubes with the fluid supply or fluid return plenums. The GE design involved passing the glass tubes through an inner wall of the structure, which necessitates complicated fabrication procedures.
It would be desirable, therefore, to develop an energy-exchange system for use with aircraft surfaces having only a single outer skin. It would also be desirable to develop an energy-exchange system in which small-diameter tubes formed to have small radii of curvature and being relatively impervious to stress corrosion and contamination corrosion could be utilized. Additionally, it would be desirable if an energy-exchange system could be developed utilizing a fluid which may be selected from either liquids or gases, in order to maximize the efficiency of the energy-exchange process.